Generally, one type of deployable solar array structure includes one or more panels that each support one or more solar cells and a deployment structure for transitioning the panels from a stowed/undeployed state to an unstowed/deployed state. In the stowed state, the panel/panels is/are typically disposed in a predefined space and orientation such that the solar cells associated with the panel/panels are either not functional or marginally functional. For example, a deployable solar array structure that includes several panels and is associated with spacecraft may have a stowed state in which the panels are disposed in a stack that is situated adjacent to the side of the spacecraft. In such a stowed state, most and potentially all of the solar cells supported by the panels are either non-functional or only marginally functional. In the unstowed/deployed state, the panel/panels is/are in an orientation/orientations such that the solar cell/cells can become functional to the extent required by the particular application. For instance, if the deployable solar array structure has a single panel that is disposed adjacent to a spacecraft, deployment of the panel may involve translating and/or rotating the panel relative to the spacecraft so that the solar cell/cells associated with the panel can be used to produce the power needed by the spacecraft. In the case of a deployable solar array structure comprised of multiple panels each associated with a “petal” structure, the petals are transitioned from the stowed state in which the petal are stacked one on top of another to a deployed state in which one or more of the petals is/are rotated so that each petal occupies a distinct radial space that exposes the solar cell/cells associated with the petal so that the cell/cells can be used to satisfy the power requirement of the spacecraft. In the situation in which a deployable solar array structure comprised of multiple panels that are connected to one another such that the panels can be “accordion” folded to form a stack, the stack of panels is unfolded such that the panels are substantially coplanar with one another and the cell/cells associated with each panel can be used to satisfy the power requirements of the spacecraft.
In many applications, deployable solar array structures that include one or more panels that each support one or more solar cells and a deployment structure for transitioning the panels from a stowed/undeployed state to an unstowed/deployed state support the panel or panels in a cantilever manner. For example, in the case of a deployable solar array structure with a single rectangular panel having two end edges and two side edges that each extend between the two end edges, one of the end edges of the panel is anchored to a support structure. The other end edge and substantially all of the structure between the two end edges is not supported. The cantilever approach avoids the need for other bracing extending between the support structure and the panel. However, the cantilever approach also limits the distance that the panel can extend away from the support structure and, as such, the area of a panel that can support a solar cell or cells. More specifically, as the distance between the supported end of the panel and the free end of the panel increases for a panel made of a given material and having given dimensions, the panel will increasingly bend or deform. This bending or deformation can be significant enough that the solar cell or cells associated with the panel cannot all be positioned to provide the needed power or the panel exceeds its stress limit and fails.
One approach to increasing the distance that a cantilevered panel or group of cantilevered panels can extend from a support structure and the area of the panel or panels that can support a solar cell/cells is to provide a panel that has a high moment of inertia and stiffness when the panel is in the deployed state. An example of this approach is set forth in U.S. Pat. No. 6,147,294 (the '294 patent). In the '294 patent, a cantilevered solar array wing that has a D-shaped cross-section in a deployed state is disclosed. Apparently, the D-shape yields the needed high moment of inertia and stiffness for the wing to extend a substantial distance from a box that supports the wing in a stowed state and is, in use, somehow associated with a spacecraft. The wing comprises five panels. Each of the panels includes an upper surface structure with four corners, a bottom surface structure with four corners, a solar cell supported by the bottom surface, a 180° strain energy type hinge extending between each of the four pairs of corners associated with the upper and bottom surface structures, and three panel-to-panel hinges for connecting each of the five panels to an adjacent panel. When the panel is in the stowed state, the upper surface structure and bottom surface structure are each flat and the 180° strain energy type hinges are each in a strained state with significant stored potential energy. In transitioning from the stowed state to the deployed state, the energy stored in the 180° strain hinges is used to bow the upper surface structure. Due to the stiffness of the bottom surface structure that supports the solar cell, the bottom surface structure remains flat during the transition of the panel between the stowed and deployed states. The bowed upper surface and the flat bottom surface define the D-shaped cross-section that increases the moment of inertia and stiffness of the panel. The panel-to-panel hinges allow several of the panels to be connected in series to realize the solar array wing that has a desired moment of inertia and stiffness that can support solar cells disposed across the bottom surfaces of the panels in a planar fashion.